This invention relates to a blade for a turbine engine. Specifically, the invention relates to a fan blade for a gas turbofan engine.
FIG. 1 provides a cross-sectional view of a gas turbofan engine 50 in a nacelle N. Briefly, air enters an inlet 51 in the nacelle N. A fan section 53 compresses the air entering the inlet 51. The fan section 53 also splits the air into a primary, or core, engine flow C and a secondary, or bypass, flow B. From this point, these flows will travel different paths through the engine.
The core engine flow C enters a compressor section of the engine. Typically, the compressor section includes a low pressure compressor 55 and a high pressure compressor 57. The compressor section increases the pressure of the air to aid in the combustion cycle.
The compressed core engine flow C then enters a diffuser/combustor section 59. The diffuser decreases the velocity of the core engine flow C and further increases pressure. The combustor section 59 mixes the core engine flow C with fuel (not shown) and combusts the mixture.
The gases from the combustor section 59 then enter a turbine section. Typically, the turbine section includes a high pressure turbine 61 connected to the high pressure compressor 57 and a low pressure turbine 63 connected to the low pressure compressor 55 and fan.
After driving the high pressure turbine 61 and the low pressure turbine 63, the core engine flow C exits the engine 50 through a nozzle 65. The core engine flow C through the nozzle 65 produces thrust.
The bypass flow B avoids the core engine. Instead, the bypass flow B travels around the core engine by following the fan section 53 and exiting through a nozzle 67. The bypass flow B through the nozzle 67 also produces thrust. The thrust produced by the bypass flow B in high bypass ratio turbofans can account for a significant portion (e.g. 75 percent) of total engine thrust.
As thrust requirements increase, designers typically increase the diameter of the engine 50. While producing greater thrust, the larger engine adds weight to the aircraft. A portion of the weight increase occurs directly within the engine. For example, the larger engine has larger and heavier fan blades that require, for example, heavier disks, bearings and supports. A portion of the weight increase also occurs indirectly. For instance, larger fan blades require a stronger containment structure to absorb a blade loss. Also, a larger engine requires a stronger pylon on the aircraft and larger struts, flanges, supports and mounts on the nacelle.
Thus, a need exists for keeping weight increases to a minimum. In fact, a preference exists for reducing weight whenever possible.
It is an object of the present invention to reduce engine weight.
It is a further object of the present invention to reduce fan blade weight.
It is a further object of the present invention to reduce the size of the retention structure that secures the fan blade to the disk.
It is a further object of the present invention to reduce the size of the disk.
It is a further object of the present invention to increase mass flow through the fan while keeping engine diameter constant.
It is a further object of the present invention to reduce blade length while keeping mass flow through the fan constant.
It is a further object of the present invention to decrease the kinetic energy of the blade during a blade loss event.
It is a further object of the present invention to reduce the size of the containment structure used to confine a released blade.
It is a further object of the present invention to decrease the unbalanced load on the rotor after a blade loss event.
It is a further object of the present invention to reduce the structural requirements of the engine and aircraft, such as the size of the engine cases, struts, flanges, supports, mounts and engine pylons.
These and other objects of the present invention are achieved in one aspect by a blade for a turbine engine having a centerline. The blade comprises: a root section extending at an angle relative to the centerline; and an airfoil section extending from the root section. The root section is directly adjacent said airfoil section
These and other objects of the present invention are achieved in another aspect by a blade for a turbine engine having an axial direction. The blade comprises: an axially oriented root section; and an airfoil section extending from the root section. The blade does not have a neck between the root section and the airfoil section.
These and other objects of the present invention are achieved in another aspect by a rotor assembly for a turbine engine having an axial direction. The rotor assembly includes: a disk having a plurality of axially oriented grooves; and a plurality of neckless blades. Each blade has a root section with a continuous enlarged head for placement within a corresponding groove.
These and other objects of the present invention are achieved in another aspect by a turbofan engine having a flow path. The engine comprises: a fan section; a compressor section; a burner section; a turbine section; and an exhaust section. The fan section includes a disk and a plurality of blades secured thereto. The outer surface of the disk and a portion of the blades define an inner boundary of the flow path.